A gas turbine engine component is provided with at least one coolant system embedded within an airfoil portion, which coolant system has at least one exit and means for preventing deposits from interfering with a flow of cooling fluid from the at least one exit.
The design of an advanced high pressure turbine component, such as a high pressure turbine vane, requires that the airfoil portion of the component be cooled with a series of highly convective coolant systems embedded in an airfoil wall. Due to the configuration of the coolant system exits, deposits have a high propensity to accumulate there. As a result, the exit planes have reduced cooling film traces due to exit plugging. When this happens, film cooling of the airfoil wall becomes affected negatively to the point where the local cooling effectiveness is affected adversely. Note that the overall cooling effectiveness is a form of the dimensionless metal temperature ratio for the airfoil. In general, the overall cooling effectiveness of this type of high pressure turbine component is close to 0.7 (unity being the maximum value), and due to film exit deposits, the cooling effectiveness can be lowered to values below 0.2. As a result, the local life capability of the part becomes very limited. Consequences of this limitation result in premature oxidation, erosion and thermal-mechanical fatigue cracking. It is therefore necessary to alleviate this problem.